Modern structures are increasingly manufactured using “composite” materials, where reinforcing materials, such as fibers, are embedded in a polymer matrix, such as a thermoset resin. Composite materials are particularly important where an important feature of the structure is the ratio of strength to weight. Aircraft components are a primary example of structures where the strength to weight ratio is a primary consideration.
One process that is commonly used to form structures from composite materials is a lay-up process. In the lay-up process, the reinforcing materials layers of reinforcing material are laid in a mold, either by hand or by machine. The polymer matrix is then introduced into the mold, filling the voids between the reinforcing materials. The reinforcing materials may be pre-impregnated with resin prior to placement in the mold. The materials are then cured, often by using an autoclave. Alternatively, the materials can be partially cured and then placed in a mold or on a mandrel for forming into a desired shape.
More complex composite structures are commonly built with discrete components. For example, referring to FIG. 1, aircraft structures are often designed with a core component 90, which may comprise a composite material or stiffening support such as a honeycomb structure, and a spar 45, arranged between two skin components 70, 71. In this example, the aircraft structure is formed by placing a partially cured or uncured first skin component 71 in a bond assembly jig, which has a support plate 5 having the desired shape for one surface of the aircraft structure. Additional layers of core component 90 are added, followed by a second cured or uncured skin component 70. A caul component may then be placed over the second skin component, having features that conforms the second skin component into a desired shape. At one end of the structure, the bond assembly jig may comprise a nose block 10 which holds the cured spar 45 in relation to the core component 90 and skin components 70, 71. The bond assembly jig or tool is then typically enclosed in an autoclave 80, which exerts heat and pressure, (represented by arrows in FIG. 1), onto the assembly in order to cure the composite materials.
When using a bond assembly jig or tool configured with a block member, a common problem occurs in which the pins 12 holding the block member tend to bind due to the lateral force exerted by the autoclave on the block member. Not only does this problem create difficulty for disassembling and reassembling the bond assembly jig. Further, the one-piece block member 10 is rigid, often leaving the adjacent spar 45 and first skin component 71 with excessive porosity or voids.
Accordingly, there is a need for an improved bond assembly jig and method which will prevent the block member pins or fasteners from binding and which will further allow the block member to float in at least one direction.